The present invention relates to systems for cooling detonation engines, and more particularly to vapor cooling systems for cooling detonation engines.
Known gas turbine engines have utilized superalloys, thermal barrier coatings (TBCs), and film cooling schemes in order to provide combustion chamber structures that can operate efficiently at high temperatures and pressure while still maintaining a relatively long component lifespan. In order to improve engine efficiency, it has been desired to develop engines that utilize detonation in addition to or instead of slower speed, non-detonative combustion. Utilization of detonation schemes (e.g., pulse detonation or continuous detonation) takes advantage of the thermodynamic benefits of fuel burn recovery, to increase engine fuel efficiency.
However, detonation engines present a number of difficulties in providing cooling to engine components in a manner that is reliable and effective. In conventional non-detonation engines, maximum temperatures are typically no more than about 1,649° C. (3,000° F.). With detonation engines, maximum temperatures can be as high as about 2,538° C. (4,600° F.). At those higher temperatures present in detonation engines, TBCs applied to superalloys may not provide sufficient thermal protection or be sufficiently durable. It is desired to maintain metal alloy temperatures below about 1,093° C. (2,000° F.) during operation. Moreover, sections of detonation engines, especially sections of pulse detonation engines, can be subject to sudden pressure spikes and turbulence as a natural consequence of the detonation process. The pressure spikes and turbulence may cause disruptive backflow or aspiration through film cooling holes, that is, acute pressure spikes or turbulence may cause hot flowpath gases to flow “backwards” into cooling holes or otherwise disrupt the film cooling. Such backflow or aspiration may make film cooling unsuitable for use with detonation engines.